In current gas turbine engines, the provision of cooling air to components downstream of the combustor is commonplace. The cooling air can allow components to operate in allowable material temperature ranges. In the case of turbine blade cooling, the air is bled from the compressor stage using a series of feed pipes and valves.
In turboprop or open rotor gas turbine engines, a row of rows of propeller blades are mounted forward of the engine or rearward of the engine. Rearward mounting can provide aerodynamic and operational advantages over forward mounting, but a problem with the arrangement is that hot exhaust gases from the engine have to traverse the propeller blades on being expelled from the engine.
In one configuration, the exhaust gases are expelled from the engine such that they impinge on the root regions of the propeller blades. As the exhaust gas can reach temperatures of about 750 K, the blades typically require cooling at the root and hub to avoid material fatigue and failure, as well as to avoid damage to the lubrication system of support bearings.
One option for cooling the blades is to supply a flow of air bled from the compressor of the engine to internal cooling networks in the root regions of the blades. However, air bled from the compressor has a detrimental effect on compressor performance. Further, the network of feed pipes and valves which the supply requires adds weight to the engine. Also, the air has to cross a rotating boundary to reach the rear rotors.
US 2009/0202357 proposes having an annular cooling flow nozzle radially inward of an annular exhaust nozzle. Exhaust gas from the exhaust nozzle impinges on a row of propeller blades. Cooling air, which may be bled from the compressor, is directed by the cooling flow nozzle essentially under the annular exhaust gas flow to enter each blade at an intake at the blade root section.
There is a need for improvements to known cooling arrangements for open rotor gas turbine engines.